Gas turbine engine with power density range

ABSTRACT

A gas turbine engine includes a propulsor section including a propulsor, a compressor section including a low pressure compressor and a high pressure compressor, a geared architecture, a turbine section including a low pressure turbine and a high pressure turbine, and a power density of greater than or equal to 4.75 and less than or equal to 5.5 lbf/in 3 , wherein the power density is a ratio of a thrust provided by the engine to a volume of the turbine section.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.17/037,916, filed Sep. 30, 2020, which is a continuation of U.S. patentapplication Ser. No. 16/186,811, filed Nov. 12, 2018, which is acontinuation of U.S. patent application Ser. No. 14/593,056, filed Jan.9, 2015, which is a continuation-in-part of U.S. patent application Ser.No. 13/446,312, filed Apr. 13, 2012, which claims priority to U.S.Provisional Application No. 61/619,111, filed Apr. 2, 2012.

BACKGROUND OF THE INVENTION

This application relates to a geared turbofan gas turbine engine,wherein the low and high pressure spools rotate in the same directionrelative to each other.

Gas turbine engines are known, and typically include a fan deliveringair into a compressor section, and outwardly as bypass air to providepropulsion. The air in the compressor is delivered into a combustionsection where it is mixed with fuel and burned. Products of thiscombustion pass downstream over turbine rotors, driving them to rotate.Typically there are low and high pressure compressors, and low and highpressure turbines.

The high pressure turbine typically drives the high pressure compressoras a high spool, and the low pressure turbine drives the low pressurecompressor and the fan. Historically, the fan and low pressurecompressor were driven at a common speed.

More recently, a gear reduction has been provided on the low pressurespool such that the fan and low pressure compressor can rotate atdifferent speeds. It desirable to have more efficient engines that havemore compact turbines to limit efficiency loses.

SUMMARY

In a featured embodiment, a gas turbine engine turbine comprises a highpressure turbine configured to rotate with a high pressure compressor asa high pressure spool in a first direction about a central axis. A lowpressure turbine is configured to rotate with a low pressure compressoras a low pressure spool in the first direction about the central axis. Apower density is greater than or equal to about 1.5 and less than orequal to about 5.5 lbf/in³. A fan is connected to the low pressure spoolvia a speed changing mechanism and will rotate in a second directionopposed to the first direction.

In another embodiment according to the previous embodiment, the powerdensity is greater than or equal to about 2.0.

In another embodiment according to any of the previous embodiments, thepower density is greater than or equal to about 4.0.

In another embodiment according to any of the previous embodiments, thepower density thrust is calculated using a value that is sea leveltake-off, flat-rated static thrust.

In another embodiment according to any of the previous embodiments,guide vanes are positioned upstream of a first stage in the low pressureturbine to direct gases downstream of the high pressure turbine as theyapproach the low pressure turbine.

In another embodiment according to any of the previous embodiments, amid-turbine frame supports the high pressure turbine.

In another embodiment according to any of the previous embodiments, theguide vanes are positioned intermediate the mid-turbine frame and thelow pressure turbine.

In another embodiment according to any of the previous embodiments, theguide vanes are highly cambered such that the vanes direct products ofcombustion downstream of the high pressure turbine to be properlydirected when initially encountering the first stage of the low pressureturbine.

In another embodiment according to any of the previous embodiments, thefan section delivers a portion of air into a bypass duct and a portionof the air into the low pressure compressor as core flow, and has abypass ratio greater than 6.

In another embodiment according to any of the previous embodiments, thespeed changing mechanism is a gear reduction.

In another embodiment according to any of the previous embodiments, astar gear is utilized to change the direction of rotation between thefan and the low pressure spool.

In another embodiment according to any of the previous embodiments, thestar gear arrangement has a gear ratio above 2.3:1, meaning that the lowpressure spool turns at least or equal to about 2.3 times as fast as thefan.

In another embodiment according to any of the previous embodiments, thespeed changing mechanism is a gear reduction.

In another embodiment according to any of the previous embodiments, astar gear is utilized to change the direction of rotation between thefan and the low pressure spool.

In another embodiment according to any of the previous embodiments, thestar gear arrangement has a gear ratio above 2.3:1, meaning that the lowpressure spool turns at least or equal to about 2.3 times as fast as thefan.

In another featured embodiment, a gas turbine engine turbine comprises ahigh pressure turbine configured to rotate with a high pressurecompressor as a high pressure spool in a first direction about a centralaxis. A low pressure turbine is configured to rotate in the firstdirection about the central axis. A power density is greater than orequal to about 4.0. A fan is connected to the low pressure turbine via agear reduction and will rotate in a second direction opposed to thefirst direction.

In another embodiment according to the previous embodiment, the powerdensity is a ratio of a thrust provided by the engine to a volume of aturbine section including both the high pressure turbine and the lowpressure turbine. The thrust is sea level take-off, flat-rated staticthrust.

In another embodiment according to any of the previous embodiments, thefan section delivers a portion of air into a bypass duct and a portionof the air into the low pressure compressor as core flow, and has abypass ratio greater than 6.

In another embodiment according to any of the previous embodiments, astar gear is utilized to change the direction of rotation between thefan and the low pressure spool.

In another embodiment according to any of the previous embodiments,there is an intermediate turbine section, which drives a compressorrotor.

In another embodiment according to any of the previous embodiments, thegear reduction is positioned intermediate the fan and a compressor rotordriven by the low pressure turbine.

In another embodiment according to any of the previous embodiments, thegear reduction is positioned intermediate the low pressure turbine and acompressor rotor driven by the low pressure turbine.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 schematically shows rotational features of one type of such anengine.

FIG. 3 is a detail of the turbine section volume.

FIG. 4 shows another embodiment.

FIG. 5 shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude, for example, three-spools, an augmentor section, or a differentarrangement of sections, among other systems or features. The fansection 22 drives air along a bypass flowpath B while the compressorsection 24 drives air along a core flowpath C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a turbofan gas turbine enginein the disclosed non-limiting embodiment, it should be understood thatthe concepts described herein are not limited to use with turbofans asthe teachings may be applied to other types of turbine engines. Forpurposes of this application, the terms “low” and “high” as applied tospeed or pressure are relative terms. The “high” speed and pressurewould be higher than that associated with the “low” spools, compressorsor turbines, however, the “low” speed and/or pressure may actually be“high.”

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. The terms “high” and “low” in relation to both the speed andpressure of the components are relative to each other, and not to anabsolute value. A combustor 56 is arranged between the high pressurecompressor 52 and the high pressure turbine 54. A mid-turbine frame 57of the engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The mid-turbineframe 57 further supports bearing systems 38 in the turbine section 28.The inner shaft 40 and the outer shaft 50 are concentric and rotate viabearing systems 38 about the engine central longitudinal axis A which iscollinear with their longitudinal axes.

The core airflow C is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path and act as inlet stator vanes to turn theflow to properly feed the first blades of the low pressure turbine. Theturbines 46, 54 rotationally drive the respective low speed spool 30 andhigh speed spool 32 in response to the expansion.

The engine 20 has bypass airflow B, and in one example is a high-bypassgeared aircraft engine. The bypass ratio may be defined as the amount ofair delivered into the bypass duct divided by the amount delivered intothe core flow. In a further example, the engine 20 bypass ratio isgreater than about six (6), with an example embodiment being greaterthan ten (10), the geared architecture 48 is an epicyclic gear train,such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3 and the low pressure turbine46 has a pressure ratio that is greater than about 5. In one disclosedembodiment, the engine 20 bypass ratio is greater than about ten (10:1),the fan diameter is significantly larger than that of the low pressurecompressor 44, and the low pressure turbine 46 and the low pressureturbine has a pressure ratio that is greater than about 5:1. Lowpressure turbine 46 pressure ratio is the total pressure measured priorto inlet of low pressure turbine 46 as related to the pressure at theoutlet of the low pressure turbine 46 prior to an exhaust nozzle. Thegeared architecture 48 may be a star gear arrangement such that the fanwill rotate in a different direction than the low spool. It should beunderstood, however, that the above parameters are only exemplary of oneembodiment of a geared architecture engine and that the presentinvention is applicable to other gas turbine engines including directdrive turbofans.

A greatest amount of thrust is provided by the bypass flow B due to thehigh bypass ratio. The fan section 22 of the engine 20 is designed for aparticular flight condition—typically cruise at about 0.8 Mach and about35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with theengine at its best fuel consumption—also known as “bucket cruise ThrustSpecific Fuel Consumption (‘TSFC’)”—is the industry standard parameterof lbm of fuel being burned per hour divided by lbf of thrust the engineproduces at that minimum point. “Low fan pressure ratio” is the pressureratio across the fan blade alone, before the Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram degR)/518.7){circumflex over ( )}0.5]. The “Low corrected fan tip speed” asdisclosed herein according to one non-limiting embodiment is less thanabout 1150 ft/second at the same cruise point.

FIG. 2 shows detail of an engine 120, which may generally have thefeatures of engine 20 of FIG. 1 . A fan 122 is positioned upstream of alow pressure compressor 124, which is upstream of a high pressurecompressor 126. A combustor 128 is positioned downstream of the highpressure compressor 126. A mid-turbine frame 142 may be positioned at adownstream end of the high pressure turbine 130, and supports a bearing138, shown schematically, to support the aft end of the high pressureturbine 130, and a high pressure spool 132. A low pressure turbine 134is positioned downstream of a mid-turbine frame 142. A low spool 136,driven by the low pressure turbine 134, drives the low pressurecompressor 124. The speed change mechanism 48 causes the fan 122 torotate at a different speed than the low pressure compressor 134. Inembodiments of this invention, the speed input to output ratio for thespeed change mechanism is above or equal to 2.3:1, and up to less thanor equal to 13:1. The gear also causes fan 122 to rotate in an opposeddirection relative to the low pressure compressor 124. As mentionedabove, a star gear arrangement may be utilized to cause the fan 122 torotate in the opposed direction (“+”) relative to the low pressurecompressor 124. In this embodiment the fan generally has less than 26blades, and the low pressure turbine has at least three stages, and upto six stages. The high pressure turbine generally has one or two stagesas shown.

In this particular embodiment, the low pressure compressor 124 and thelow pressure turbine 134 rotate in one direction (“−”) and the highpressure turbine 130, the high pressure compressor 126, rotate in thesame direction (“−”).

A strut 140 is shown between the low pressure compressor 124 and thehigh pressure compressor 126. The strut 140 spans the gas path, and hasan airfoil shape, or at least a streamline shape. The combination of ablade at the exit of the low pressure compressor 124, the strut 140, anda variable vane, and then the first blade of the high pressurecompressor 126 is generally encompassed within the structure illustratedas the strut 140.

Since the compressor sections 124 and 126 rotate in the same direction,the several airfoils illustrated as the element 140 are required to doless turning of the air flow.

As will be explained below, since the turbine section is provided with ahighly cambered vane, there is less turning required between the twoturbine sections. Since the compressor is forcing flow with an adversepressure gradient, and whereas the turbine has a favorable pressuregradient, this overall engine architecture is benefited by theillustrated combination.

Highly cambered inlet guide vanes 143 are positioned in a locationintermediate the mid-turbine frame 142 and the most upstream rotor inthe low pressure turbine 134. The vanes 143 must properly direct theproducts of combustion downstream of the high pressure turbine 130 asthey approach the first rotor of the low pressure turbine 134. It isdesirable for reducing the overall size of the low pressure turbine thatthe flow be properly directed when it initially encounters the firststage of the low pressure turbine section.

The above features achieve a more compact turbine section volumerelative to the prior art, including both the high and low pressureturbines. A range of materials can be selected. As one example, byvarying the materials for forming the low pressure turbine, the volumecan be reduced through the use of more expensive and more exoticengineered materials, or alternatively, lower priced materials can beutilized. In three exemplary embodiments the first rotating blade of theLow Pressure Turbine can be a directionally solidified casting blade, asingle crystal casting blade or a hollow, internally cooled blade. Allthree embodiments will change the turbine volume to be dramaticallysmaller than the prior art by increasing low pressure turbine speed. Inaddition, high efficiency blade cooling may be utilized to furtherresult in a more compact turbine section.

Due to the compact turbine section, a power density, which may bedefined as thrust in pounds force produced divided by the volume of theentire turbine section, may be optimized. The volume of the turbinesection may be defined by an inlet of a first turbine vane in the highpressure turbine to the exit of the last rotating airfoil in the lowpressure turbine, and may be expressed in cubic inches. The staticthrust at the engine's flat rated Sea Level Takeoff condition divided bya turbine section volume is defined as power density. The sea leveltake-off flat-rated static thrust may be defined in lbs force, while thevolume may be the volume from the annular inlet of the first turbinevane in the high pressure turbine to the annular exit of the downstreamend of the last rotor section in the low pressure turbine. The maximumthrust may be Sea Level Takeoff Thrust “SLTO thrust” which is commonlydefined as the flat-rated static thrust produced by the turbofan atsea-level.

The volume V of the turbine section may be best understood from FIG. 3 .As shown, the frame 142 and vane 143 are intermediate the high pressureturbine section 130, and the low pressure turbine section 134. Thevolume V is illustrated by dashed line, and extends from an innerperiphery I to an outer periphery O. The inner periphery is somewhatdefined by the flowpath of the rotors, but also by the inner platformflow paths of vanes. The outer periphery is defined by the stator vanesand outer air seal structures along the flowpath. The volume extendsfrom a most upstream end of the vane 400, typically its leading edge,and to the most downstream edge 401 of the last rotating airfoil in thelow pressure turbine section 134. Typically this will be the trailingedge of that airfoil.

The power density in the disclosed gas turbine engine is much higherthan in the prior art. Eight exemplary engines are shown below whichincorporate turbine sections and overall engine drive systems andarchitectures as set forth in this application, and can be found inTable I as follows:

TABLE 1 Thrust Turbine Thrust/turbine SLTO section volume section volumeEngine (lbf) from the Inlet (lbf/in³) 1 17,000 3,859 4.41 2 23,300 5,3304.37 3 29,500 6,745 4.37 4 33,000 6,745 4.84 5 96,500 31,086 3.10 696,500 62,172 1.55 7 96,500 46,629 2.07 8 37,098 6,745 5.50

Thus, in embodiments, the power density would be greater than or equalto about 1.5 lbf/in³. More narrowly, the power density would be greaterthan or equal to about 2.0 lbf/in³.

Even more narrowly, the power density would be greater than or equal toabout 3.0 lbf/in³.

More narrowly, the power density is greater than or equal to about 4.0lbf/in³. More narrowly, the power density is greater than or equal toabout 4.5 lbf/in³. Even more narrowly, the power density is greater thanor equal to about 4.75 lbf/in³. Even more narrowly, the power density isgreater than or equal to about 5.0 lbf/in³.

Also, in embodiments, the power density is less than or equal to about5.5 lbf/in³.

While certain prior engines have had power densities greater than 1.5,and even greater than 3.2, such engines have been direct drive enginesand not associated with a gear reduction. In particular, the powerdensity of an engine known as PW4090 was about 1.92 lbf/in³, while thepower density of an engine known as V2500 had a power density of 3.27lbf/in³.

Engines made with the disclosed architecture, and including turbinesections as set forth in this application, and with modifications comingfrom the scope of the claims in this application, thus provide very highefficient operation, and increased fuel efficiency and lightweightrelative to their trust capability.

FIG. 4 shows an embodiment 200, wherein there is a fan drive turbine 208driving a shaft 206 to in turn drive a fan rotor 202. A gear reduction204 may be positioned between the fan drive turbine 208 and the fanrotor 202. This gear reduction 204 may be structured and operate likethe gear reduction disclosed above. A compressor rotor 210 is driven byan intermediate pressure turbine 212, and a second stage compressorrotor 214 is driven by a turbine rotor 216. A combustion section 218 ispositioned intermediate the compressor rotor 214 and the turbine section216.

FIG. 5 shows yet another embodiment 300 wherein a fan rotor 302 and afirst stage compressor 304 rotate at a common speed. The gear reduction306 (which may be structured as disclosed above) is intermediate thecompressor rotor 304 and a shaft 308 which is driven by a low pressureturbine section.

The FIG. 4 or 5 engines may be utilized with the density featuresdisclosed above.

Although an embodiment of this invention has been disclosed, a person ofordinary skill in this art would recognize that certain modificationswould come within the scope of this application. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

What is claimed is:
 1. A gas turbine engine comprising: a propulsorsection including a propulsor having a plurality of blades; a compressorsection including a low pressure compressor and a high pressurecompressor; a geared architecture; a turbine section including a lowpressure turbine and a high pressure turbine, the low pressure turbineincluding three stages, the high pressure turbine including two stages,the high pressure turbine rotatable with the high pressure compressor asa high pressure spool about an engine longitudinal axis, and the lowpressure turbine rotatable as a low pressure spool about the enginelongitudinal axis, the propulsor driven by the low pressure turbinethrough the geared architecture, and the low pressure compressor havinga greater number of stages than the high pressure turbine; and a powerdensity of greater than or equal to 4.75 and less than or equal to 5.5lbf/in³, wherein the power density is a ratio of a thrust provided bythe engine to a volume of the turbine section, the thrust is sea leveltake-off, flat-rated static thrust, and the volume of the turbinesection is defined by an inlet of a first turbine vane in the highpressure turbine to an exit of a last rotating airfoil in the lowpressure turbine.
 2. The gas turbine engine as recited in claim 1,wherein the geared architecture is an epicyclic gear train.
 3. The gasturbine engine as recited in claim 2, wherein a gear ratio of the gearedarchitecture is greater than 2.3, and the propulsor has 26 or fewerblades.
 4. The gas turbine engine as recited in claim 3, wherein: thelow pressure turbine includes no more than six stages; and the lowpressure turbine includes an inlet, an outlet and a pressure ratio ofgreater than 5, the pressure ratio being pressure measured prior to theinlet as related to pressure at the outlet prior to an exhaust nozzle.5. The gas turbine engine as recited in claim 4, wherein a compressorrotor of the compressor section is driven by the low pressure turbinethrough the geared architecture.
 6. The gas turbine engine as recited inclaim 4, wherein the low pressure compressor is rotatable with the lowpressure turbine as the low pressure spool such that the low pressureturbine drives both the low pressure compressor and the gearedarchitecture.
 7. The gas turbine engine as recited in claim 6, whereinthe high pressure compressor includes eight stages.
 8. The gas turbineengine as recited in claim 7, wherein the turbine section includes amid-turbine frame arranged between the high pressure turbine and the lowpressure turbine with respect to the engine longitudinal axis, themid-turbine frame supports a bearing, and the mid-turbine frame includesairfoils in a core airflow path.
 9. The gas turbine engine as recited inclaim 8, wherein the power density is between 4.84 lbf/in³ and 5.5lbf/in³.
 10. The gas turbine engine as recited in claim 9, wherein thehigh pressure spool and the low pressure spool are rotatable in a firstdirection about the engine longitudinal axis, and the propulsor isrotatable about the engine longitudinal axis in a second directionopposed to the first direction.
 11. The gas turbine engine as recited inclaim 7, wherein the geared architecture is a star gear arrangement, thelow pressure spool is rotatable in a first direction about the enginelongitudinal axis, and the propulsor is rotatable in a second directionopposed to the first direction.
 12. The gas turbine engine as recited inclaim 11, wherein: the propulsor section is a fan section, the propulsoris a fan, and an outer housing surrounds the fan to define a bypassduct; the fan section delivers a portion of air into the compressorsection, and a portion of air into the bypass duct, and a bypass ratio,which is defined as a volume of air passing to the bypass duct comparedto a volume of air passing into the compressor section, is greater than10; and the fan has a low fan pressure ratio of less than 1.45 acrossthe fan blades alone at cruise at 0.8 Mach and 35,000 feet.
 13. The gasturbine engine as recited in claim 12, wherein the power density is atleast 4.84 lbf/in³.
 14. The gas turbine engine as recited in claim 13,wherein the high pressure compressor and the low pressure compressorhave an equal number of stages.
 15. The gas turbine engine as recited inclaim 13, wherein the low pressure compressor includes a greater numberof stages than the low pressure turbine.
 16. The gas turbine engine asrecited in claim 13, wherein the turbine section includes a mid-turbineframe arranged between the high pressure turbine and the low pressureturbine with respect to the engine longitudinal axis, the mid-turbineframe supports a bearing arranged to support the high pressure turbine,and the mid-turbine frame includes airfoils in a core airflow path. 17.The gas turbine engine as recited in claim 16, wherein the power densityis greater than or equal to 5.0.
 18. The gas turbine engine as recitedin claim 7, wherein the geared architecture is a planetary gear system.19. The gas turbine engine as recited in claim 18, wherein: thepropulsor section is a fan section, the propulsor is a fan, and an outerhousing surrounds the fan to define a bypass duct; the fan sectiondelivers a portion of air into the compressor section, and a portion ofair into the bypass duct, and a bypass ratio, which is defined as avolume of air passing to the bypass duct compared to a volume of airpassing into the compressor section, is greater than 10; and the fan hasa low fan pressure ratio of less than 1.45 across the fan blades aloneat cruise at 0.8 Mach and 35,000 feet.
 20. The gas turbine engine asrecited in claim 19, wherein the power density is at least 4.84 lbf/in³.21. The gas turbine engine as recited in claim 20, wherein the highpressure compressor and the low pressure compressor have an equal numberof stages.
 22. The gas turbine engine as recited in claim 20, whereinthe low pressure compressor includes a greater number of stages than thelow pressure turbine.
 23. The gas turbine engine as recited in claim 20,wherein the turbine section includes a mid-turbine frame arrangedbetween the high pressure turbine and the low pressure turbine withrespect to the engine longitudinal axis, the mid-turbine frame supportsa bearing arranged to support the high pressure turbine, and themid-turbine frame includes airfoils in a core airflow path.
 24. The gasturbine engine as recited in claim 23, wherein the power density isgreater than or equal to 5.0.
 25. A gas turbine engine comprising: apropulsor section including a propulsor having a plurality of blades; acompressor section including a low pressure compressor and a highpressure compressor; a geared architecture; a turbine section includinga low pressure turbine and a high pressure turbine, the low pressureturbine including three stages, the high pressure turbine including twostages, the high pressure turbine rotatable with the high pressurecompressor as a high pressure spool about an engine longitudinal axis,the low pressure turbine rotatable as a low pressure spool about theengine longitudinal axis, the propulsor driven by the low pressureturbine through the geared architecture, the turbine section including amid-turbine frame arranged between the high pressure turbine and the lowpressure turbine with respect to the engine longitudinal axis, and themid-turbine frame supporting a bearing arranged to support the highpressure turbine; and a power density of greater than or equal to 4.75and less than or equal to 5.5 lbf/in³, wherein the power density is aratio of a thrust provided by the engine to a volume of the turbinesection, and the thrust is sea level take-off, flat-rated static thrust,and the volume of the turbine section is defined by an inlet of a firstturbine vane in the high pressure turbine to an exit of a last rotatingairfoil in the low pressure turbine.
 26. The gas turbine engine asrecited in claim 25, wherein the geared architecture is an epicyclicgear train.
 27. The gas turbine engine as recited in claim 26, wherein:the low pressure turbine includes no more than six stages; and the lowpressure turbine includes an inlet, an outlet and a pressure ratio ofgreater than 5, the pressure ratio being pressure measured prior to theinlet as related to pressure at the outlet prior to an exhaust nozzle.28. The gas turbine engine as recited in claim 27, wherein: thepropulsor section is a fan section, the propulsor is a fan, and an outerhousing surrounds the fan to define a bypass duct; the fan sectiondelivers a portion of air into the compressor section, and a portion ofair into the bypass duct, and a bypass ratio, which is defined as avolume of air passing to the bypass duct compared to a volume of airpassing into the compressor section, is greater than 10; and the fan hasa low fan pressure ratio of less than 1.45 across the fan blades aloneat cruise at 0.8 Mach and 35,000 feet.
 29. The gas turbine engine asrecited in claim 28, wherein the geared architecture is a star geararrangement, the low pressure spool is rotatable in a first directionabout the engine longitudinal axis, and the propulsor is rotatable in asecond direction opposed to the first direction.
 30. The gas turbineengine as recited in claim 28, wherein the geared architecture is aplanetary gear system.